91
NASA JSC Lunar Surface Concept Study Lunar Energy Storage NNJ08TA84C U.S. Chamber of Commerce Programmatic Workshop NASA JSC Lunar Surface Concept Study Lunar Energy Storage NNJ08TA84C U.S. Chamber of Commerce Programmatic Workshop 26 February 2009 Dr. Cheng-Yi Lu Jim McClanahan Hamilton Sundstrand Energy, Space & Defense Rocketdyne 26 February 2009 Dr. Cheng-Yi Lu Jim McClanahan Hamilton Sundstrand Energy, Space & Defense Rocketdyne

NASA JSC Lunar Surface Concept Study Lunar Energy … · 6 Lunar Surface Energy Storage Assessment Process • Requirements • FOM assumptions • Literature search • Orbital models

  • Upload
    lebao

  • View
    216

  • Download
    0

Embed Size (px)

Citation preview

NASA JSC Lunar Surface Concept StudyLunar Energy Storage

NNJ08TA84CU.S. Chamber of Commerce Programmatic Workshop

NASA JSC Lunar Surface Concept StudyLunar Energy Storage

NNJ08TA84CU.S. Chamber of Commerce Programmatic Workshop

26 February 2009

Dr. Cheng-Yi LuJim McClanahan

Hamilton Sundstrand Energy, Space & Defense Rocketdyne

26 February 2009

Dr. Cheng-Yi LuJim McClanahan

Hamilton Sundstrand Energy, Space & Defense Rocketdyne

2

Agenda

• Scope/Objectives

• Trade Study Requirements, Trade Space and Figures-of-Merit

• Task I: Trade Results, Literature Search Results

• Task I: Proposed Power Beaming Concept Feasibility

• Task I: Proposed Power Beaming Technology Development

• Task II: Other Lunar Energy Storage Technologies for Comparison

• Conclusions

• Recommended Future Work

3

Scope

• There are two major tasks:– Evaluate the power beaming technologies at an appropriate orbit

and spacecraft constellation to provide lunar night power

– Provide a comparison of power beaming design with other surface energy storage technologies for lunar night power needs

4

Trade Study Requirements, Trade Space and Figures-of-Merit

5

Requirements (per Broad Agency Announcement NNJ08ZBT002)

• 2 - 5 kWe surface electrical power (user)• 100 – 2,000 kW-hr net energy storage per module• TRL 6 by 2015 - 2018 timeframe• 5 - 10 year calendar life• 10,000 - 15,000 hour operational life• 100 – 2,000 charge/discharge cycles• Ability to withstand high dust, radiation, and widely varying thermal

environment• Anywhere location on lunar surface

6

Lunar Surface Energy Storage Assessment Process

• Requirements

• FOM assumptions

• Literature search

• Orbital models– AGI

STK/Astrogator

Orbit

Halo Orbit(s) Other Orbit(s)

Power Beaming

Number of Spacecraft

Spacecraft Technology

Fast Access Spacecraft Testbed (FAST) Type

Conventional GEO Spacecraft

Beaming Technologies

Solar Flux Microwave Laser

1 2 3

Task 1: Power Beaming for Lunar Night Power Needs

7

Lunar Surface Energy Storage Assessment Process

• Requirements

• FOM assumptions

• Literature search

• Technology evaluation

Task 2: Comparison of Lunar Energy Storage Options

Li-S (Sion)LiTE*STARNano Safe

* Reference Technology: Fission Surface Power

Regenerative Fuel Battery

Separated

Flywheel

Li-IonUnitized

Lunar Surface Energy Storage*

Supercapacitor

Pressured Storage Cryogenic Storage

Thermal Storage

Thin Film

EEStor

Integrated with Propellant Tanks

SeparatedTanks

08PD-167-003

Regenerative Fuel Cells

Pressurized Storage

8

Figures-of-Merit

• Quantitative FOM’s– Mass– Launch Volume– System Efficiency– Technical Readiness Level (TRL)

• Qualitative FOM’s– Operational Effectiveness: Integration, Redundancy, Mobility, …– Development Schedule– Relative Cost

9

Task I Trade Study

10

Task 1 Conclusions

• A frozen lunar orbit (16.1 hr period) appears promising

– Provides a reasonable gap time and maximum range

10,731 km maximum range (2020)14 hour gap time (Jan to Dec 2020)15.6 hour gap time with pointing error & elevation angle constraints

• One spacecraft provides a good trade-off between cost/mass & coverage

– Two spacecraft provide redundancy• FAST spacecraft provides lower mass

and related transportation cost• Laser power beaming technology

provides smallest surface footprint.

Orbit

Halo Orbit(s) Other Orbit(s)

Power Beaming

Number of Spacecraft

Spacecraft Technology

Fast Access Spacecraft Testbed (FAST) Type

Conventional GEO Spacecraft

Beaming Technologies

Solar Flux Microwave Laser

1 2 3√

Configurations have ability to integrate telecommunications funcConfigurations have ability to integrate telecommunications function with power infrastructuretion with power infrastructure

11

Literature Search Results NASA Langley Research Center Report*

* Williams, M.D., DeYoung, R.J., Schuster, G.L., and Choi H.S., “Power Transmission by LaserBeam from Lunar-Synchronous Satellite,” NASA TM-4496, November 1993.

“ This study addresses the possibility of beaming laser power from synchronous lunar orbits…. to a manned long-range lunar rover. “

• Small receiver diameter and large laser beam (aperture) diameter due to long distance power transmission

On-orbit PV power laser beaming power to lunar rover

A ’dish-like” laser receiver mounted on lunar rover to receive and convert laser power to electric power

12

Literature Search Results NASA Glenn Research Center Report*

* Kerslake, Thomas W., “Lunar Surface-to-Surface Power Transfer,” NASA/TM-2007-215041, NASA Glenn Research Center, December 2007

“A human lunar outpost, … has potential requirements to transfer electric power up to 50-kW across the lunar surface from 0.1 to 10-km distances..“

• Trades among AC or DC power via cables, beamed radio frequencypower and beamed laser power

• Small receiver and beam aperture diameters due to short distance

Rover receiving laser power beam from Lunar LanderLunar Lander sending laser power beam to rover

13

Groundrules and Assumptions

• Space Element– 130 W/kg specific power for electrical power system (FAST)

Employs CPV cell technology

– Fixed spacecraft bus and propulsion system dry mass200 kg and 30 kg, respectively

– Yearly station-keeping ΔV budgets132 m/s for L1 and L2 halo orbits22.9 m/s for lunar polar, equatorial and 45° inclination orbits

– Laser module and dc-to-dc converter unit (DDCU) power electronicsCombined 50% efficiency and 50 W/kg composite specific power based on

– 830 nm laser wavelength– Laser diode module and DDCU power electronics specific mass

12 kg/kW and 6.3 kg/kW, respectively10 percent additional mass allocation for structure and integration

– 2 kg/m2 laser module heat rejection radiator areal density– 1,367 W/m2 solar insolation

14

Ground Rules and Assumptions Concluded

• Surface Element– Laser receiver employs thin-film (CIGS) photovoltaic technology

50 % combined efficiency, 10 year life– Based 830 nm wavelength (~ 0.7 transmission efficiency x 0.8 quantum efficiency)

1367 W/kg specific power– Based on 0.5 kg/m2 areal density and 50% efficiency

– Required infrastructure to maintain dust-free surface is not included

15

Functional Block Diagram and Energy Balance

Solar PowerCollection &

Concentration

Solar PowerConversion

Elec. PowerConversion

Laser Conversion

Power Beaming Architecture Space PowerSpace Power Element

Power Beaming Architecture Surface PowerSurface Power Element

Laser PowerTransmission/

CollectionLaser PowerConversion

Energy Storage Architecture Surface Power OnlySurface Power Only Element

Solar PowerConversion

EnergyStorage

10, 14 &

22 kWe power

generation30%

2, 3 & 5 kWeto user at night

70%* 95% (DDCU)

80%60 - 70%

~50% (Laser)

Elec. Power Conversion

95% (DDCU)

2, 3 & 5 kWeto user at night

Elec. Power Conversion

95% (DDCU)< 20% (CIGS) / 28% (5J) 95% (Li-Ion Battery)

800, 1200 &

2000 kWe-hrenergy

* Function/ Assembly Efficiency

16

Round 1: Candidate Solar Power Collection Spacecraft Orbits

A. Earth-Moon libration point L1 circular halo orbit (1 spacecraft)B. Earth-Moon libration point L2 circular halo orbit (1 spacecraft)C. Lunar circular equatorial orbit (1 spacecraft)D. Lunar circular polar orbit (1 spacecraft)E. Lunar elliptical orbit, 45° inclination, 1 to 2 planes (1 spacecraft per plane)

Note: not to scale.

17

Round 1 Candidate Orbit Definitions

• Earth-Moon libration point L1 halo orbit (1 spacecraft)– 3,700 km halo radius, near stationary with respect to lunar surface,

58,000 km distance, multiple halo orbit inclination options• Earth-Moon libration point L2 halo orbit (1 spacecraft)

– 3,700 km halo radius, near stationary with respect to lunar surface, 64,500 km distance, multiple halo orbit inclination options

• Lunar equatorial orbit (1 spacecraft)– 36 hour orbit period, 11,039 km / 0° inclination

• Lunar polar orbit (1 spacecraft)– 36 hour orbit period, 11,041 km / 0° inclination

• Lunar frozen orbit, 45° inclination, 1 plane (1 spacecraft)– 36 hour orbit period, 5,928 x 16,149 km

• Lunar frozen orbit, 45° inclination, 2 planes (1 spacecraft per plane)– 36 hour orbit period, periapsis points clocked 180 degrees apart

18

Candidate Orbit Comparison Summary

Earth-Moon libration point L1circular halo orbit (1 spacecraft)

Candidate Orbit Advantages Disadvantages

Almost always in sunlightSimpler pointing & trackingFlexible surface sites (front side)

Large surface footprint (119 m)Heavy surface receiver (27X)Much costly (> 10X)Station-keeping propellant (6X)

Almost always in sunlightSimpler pointing & trackingFlexible surface sites (far side)

Large surface footprint (132 m)Heavy surface receiver (33X)Much costly (> 12X)Station-keeping propellant (6X)

Lunar circular equatorial orbit (1 spacecraft constellation)

Lunar circular polar orbit (1 spacecraft constellation)

Lunar elliptical orbit, 45 deg inclination, 1 to 2 planes (1 to 2 spacecraft constellation)

Earth-Moon libration point L2circular halo orbit (1 spacecraft)

Small surface footprint (23 m)Minimal shadow timeMinimal station-keeping prop.

Pointing & tracking complexityLimited to equatorial sites

Small surface footprint (23 m)Minimal shadow timeMinimal station-keeping prop.

Pointing & tracking complexityLimited to single long. sites

Small surface footprint (34 m)Minimal station-keeping prop.Flexible surface sites

Pointing & tracking complexity 2X cost over single spacecraft

L1 or L2 Halo orbit is not appropriate candidate for lunar surface power beaming

19

Spacecraft Technology Comparison

FAST

Specific Power of S/C (W/kg)

Specific Power of Electric Power System (W/kg)*

Conventional S/C

< 4

< 55

40

130

* Solar power collection, power conversion, electrical power management and distribution systems, heat rejection and all supporting structures, including pointing and deployment mechanisms along with sun pointing and tracking mechanisms

FAST spacecraft with electric propulsion is a feasible candidate for lunar mission

20

Beaming Technology Trade - Footprint Comparison

Laser Solar Flux Microwave

Wavelength (μm)Beam diameter (m)Distance (km)Receiver diameter (m)Receiver area (m2)

0.8301.0

10,73122.7406

10.0751.0

10,731264.855,071

53.5341.0

10,7311,402.7

1,545,357

(5.6 GHz)

Laser-PV is the choice for lunar surface power beaming technology

21

Task I Trade Study Round 2 Additional Gs&As• 10° latitude, -45° longitude surface site• Mission time of 2020 • 0.05° laser beam pointing error• Min. of 30° elevation angle constraint (trade study results)• Energy storage architecture

– 200W-hr/kg energy storage technology– 95% charge/discharge (roundtrip) cycle efficiency

• Frozen orbit(s) trade among 8 to 48 hour circular/elliptical orbits– Result: Near-optimal elliptical orbit, 45° inclination, 16.1 hr orbit period

max. of 14 hour gap time without elevation angle constraint for year of 2020– 10,731 km maximum range

max. of 15.6 hour gap time with minimum 30° elevation angle & year of 2020– 9,977 km maximum range

• “Gap” time is defined as the time when spacecraft is not providing power to the surface site during lunar night

– S/c and surface site line-of-sight is not available or not practical or– S/c is not in direct sunlight – Surface energy storage would be required during gap time

22

Power Beaming Energy Balance with Gap Time

Solar PowerCollection &

Concentration

Solar PowerConversion

Elec. PowerConversion

Laser Conversion

Power Beaming Architecture Space PowerSpace Power Element

Power Beaming Architecture Surface PowerSurface Power Element

Laser PowerTransmission

/CollectionLaser PowerConversion

EnergyStorage

27 kWe power

generation

5 kWe to user

70%* 30% 95% (DDCU)

80%60 - 70%

~50% (Laser)

Elec. Power Conversion

95% (DDCU)

95% (Li-Ion Battery)

* Function/Assembly Efficiency

525 kW-hr

23

Max. Gap Summary – 16.1 Hour Orbit Period

8.8

5.0

0.0

13.7

11.4

6.5

14.013.1

6.9

0

2

4

6

8

10

12

14

16

1 Sat., Ellip. Orbit 2 Sat., Ellip. Orbit 3 Sat., Ellip. Orbit

Space Element Configuration

Max

imum

Gap

Tim

e (h

ours

)Jan 2020 Jan to Jun 2020 Jan to Dec 2020

STK v8.1.3 simulation results. 10° latitude, -45° longitude surface siteNo elevation angle constraints

Mission time and duration and # of s/c’s drive Gap time

24

Max. Gap Time Versus Elevation Angle Constraint

13.0

13.5

14.0

14.5

15.0

15.5

16.0

16.5

0 5 10 15 20 25 30 35 40 45 50

Elevation Angle (degrees)

Max

imum

Gap

Tim

e (h

ours

) 15.6 hours at 30° elev. angle

10° latitude, -45° longitude surface site.Elliptical orbit, 45° inclination, 16.1 hour orbit period.

25

Max. Range Versus Elevation Angle Constraint

9,600

9,700

9,800

9,900

10,000

10,100

10,200

10,300

10,400

10,500

10,600

10,700

10,800

0 5 10 15 20 25 30 35 40 45 50

Elevation Angle (degrees)

Max

imum

Ran

ge (k

m)

9,977 km at 30° elev. angle

10° latitude, -45° longitude surface site.Elliptical orbit, 45° inclination, 16.1 hour orbit period.

26

Receiver Area Variation with Pointing Error

0

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

9,000

10,000

0.00 0.05 0.10 0.15 0.20

Laser Beam Pointing Error (deg)

Luna

r Sur

face

Col

lect

or A

rea

(m2 )

(2087 m2 at 0.05° pointing error, 30 ° minimum elevation angle and 9,977 km range)

10° latitude, -45° longitude surface site.Elliptical orbit, 45° inclination, 16.1 hour orbit period.

27

Receiver Mass Variation with Pointing Error

0

500

1,000

1,500

2,000

2,500

3,000

3,500

4,000

4,500

5,000

0.00 0.05 0.10 0.15 0.20

Laser Beam Pointing Error (deg)

Luna

r Sur

face

Col

lect

or M

ass

(kg)

(1043 kg at 0.05° pointing error, 30 ° minimum elevation angle and 9,977 km range)

10° latitude, -45° longitude surface site.Elliptical orbit, 45° inclination, 16.1 hour orbit period.

28

Proposed Power Beaming Concept Feasibility

29

Laser Power Beaming Architecture Concept

30° minimum elevation angle 30° minimum elevation angle

Primary spacecraft

Redundant spacecraft

16.1 hour frozen elliptical orbit(sma = 7,472 km, e = 0.4082, i = 45°)

Surface site(10° latitude,-45° longitude)

Moon

Laser beam

Sun

30

Access and Gaps Summary – Jan through Feb 2020

0

5

10

15

01 Ja

n 202

008

Jan 2

020

15 Ja

n 202

022

Jan 2

020

29 Ja

n 202

005

Feb 20

2012

Feb 20

2019

Feb 20

2026

Feb 20

20

Date

Ope

ratio

nal T

ime

(hou

rs)

Laser Power Beaming Energy Storage

31

Lunar Surface User Power Schedule

0

5

10

15

20

25

30

35

40

45

0 5 10 15 20 25 30

Lunar Day Lunar Night

Time (days)

Use

r Pow

er (k

We)

Laser on (provides user power and to charge energy storage subsystem)Laser off (user power provided by energy storage subsystem)

32

Laser Power Beaming Architecture Description

• Space Element– 3-axis stabilized spacecraft (one primary and one redundant)– Electric propulsion for orbit and station-keeping operations

Orbit clean-up following insertion and for de-saturating CMG’s

– Solid state laser diode technology (830 nm wavelength)1 m laser beam diameter at sourceTwo-axis tracking gimbal with a pointing accuracy of 0.05°

• Surface Element– “Plug and play” CIGS photovoltaic array technology– 16 lightweight flexible panels (receiver) with “tent” structures at each end

Manual deployment facilitated by two crew membersMinimal surface pre-treatment and minimal surface support equipment

– Removal of large rocks– Simple alignment using known spacecraft ground track and surface position data

33

STK Generated 3-D Orbit Track

Elliptical orbit, 45° inclination, 16.1 hour orbit period. 10° latitude, -45° longitude surface site.

34

FAST Application: Power Beaming for Lunar Surface Power

35

STK Generated Ground Track

10° latitude, -45° longitude surface site.

Surface Site

Elliptical orbit, 45° inclination, 16.1 hour orbit period.

36

Lunar Surface Receiver

Design footprint(41.4 m x 64.1 m)

Ideal footprint + 30° elevation angle constraint(22.7 m x 45.4 m)

Ideal footprint + 0.05° laser pointing error(41.4 m)

Ideal footprint(22.7 m)

Design footprint accounts for 0.05° pointing error and 30° minimum elevation angle constraintDesign footprint accounts for 0.05° pointing error and 30° minimum elevation angle constraint

Direction of flight

Laser beam spill5 X ideal footprint area

37

Lunar Surface Receiver – 30° Min. Elevation Angle

Top View

Side View

Direction of flight

“Flat” section

16 flexible CIGS photovoltaic array modules are complemented by “tent”structures (2.5 m wide and x 70.9 m maximum length module).16 flexible CIGS photovoltaic array modules are complemented by “tent”structures (2.5 m wide and x 70.9 m maximum length module).

“Tent” section

PMAD subsystem interface

“Tent” section (to minimize laser beam spill)

Design footprint (41.4 m x 64.1 m)

38

“Tent”

• Orient the last 5 m of each end of the active photovoltaic arraymodule to an angle of 45° to form a “tent”

• The “tent” structure serves two purposes– Minimize laser beam spill– Minimize contamination of the surface receiver when surface activities

are conducted in the vicinity of the surface receiver

39

Deployed and Stowed Surface Receiver Module

2.5 m

71 m

2.5 m

Deployed Top View

Stowed Isometric View 0.1 m

1.0 m

Lunar surface receiver module stows into a reasonable sized package, 1 m x 2.5 m x 0.1 m (0.2 m3) with 71 folds.Lunar surface receiver module stows into a reasonable sized package, 1 m x 2.5 m x 0.1 m (0.2 m3) with 71 folds.

“Tent” section(at each end of module)

40

Lunar Surface Receiver Deployment Sequence

Top View

Side View

“Flat” section “Tent” section

PMAD subsystem interface

“Tent” section

Design footprint (41.4 m x 64.1 m)

Step 50 – Connect16 PV array electricalcables to PMADinterface cable

Step 49 – Connect to power grid

Lunar surface receiver may be deployed in as few as 50 steps. Each module deploys independently and can be stowed via refolding.Lunar surface receiver may be deployed in as few as 50 steps. Each module deploys independently and can be stowed via refolding.

Step 0 – Align 51 x 81 m deployment area to spacecraft ground track and grade

Graded deployment area

Direction of flight

13 24 5679 8

10 1112

13 1415

16 1718

19 2021

22 2324

25 2627

28 293034 3536

31 3233

37 3839

40 4142

43 4445

46 4748

41

Surface Element Energy Storage and Power Trade

Energy Storage Charge Total Collector E-Storage Total Comments

kW-hr kWe kWe kg kg kg

151 15 20 4,572 755 5,327

240 10 15 3,429 1,200 4,629

400 4 9 2,057 2,000 4,057

500 2 7 1,600 2,500 4,100

525 1 6 1,372 2,625 3,997 minimum charge power

Power Mass

Energy Storage Charge Total Collector E-Storage Total Comments

kW-hr kWe kWe kg kg kg

151 15 20 4,572 755 5,327

240 10 15 3,429 1,200 4,629

400 4 9 2,057 2,000 4,057

500 2 7 1,600 2,500 4,100

525 1 6 1,372 2,625 3,997 minimum charge power

Power Mass

• Total of ~ 4 MT and 3.3 m3 (stowed volume) lunar surface receiver and 525 kW-hr battery are required for a continuous 5 kWe output

• ~2/3 of the total mass results from battery

• Receiver (41 m x 64 m) can generate > 500 kWe during daytime

Power beaming configuration can reduce energy storage requirement by 70%Power beaming configuration can reduce energy storage requirement by 70%

42

Proposed Technology Development

43

Current State-of-the-Art CIGS PV Arrays

Thin Film PV Array can be deployed by two crew membersThin Film PV Array can be deployed by two crew members

44

Current State-of-the-Art CIGS PV Arrays (concluded)

• Recent NREL report of 15% efficiency for CIGS • Expect efficiency to be 20% by 2015 - 2018• Recent NREL report of 15% efficiency for CIGS • Expect efficiency to be 20% by 2015 - 2018

45

Boeing Solid-State Disk Laser Concept (June 2008)

Laser Subsystem Engineering ConceptAxisymmetric layout showing multiple disks mounted on a common substrate

– 25 kW thin-disk, solid-state tactical laser systemScalable to a 100-kW-class system based on the same architecture and technologyGiving the warfighter an ultra-precision engagement capabilityIncorporates COTS, SOTA lasers used in the automotive industry

– Have demonstrated exceedingly high reliability, supportability and maintainability

46

Laser Power Beaming Technology Roadmap

2009/2010

2011/2012

2013/2014

2015/2016

2017/2018

2019/2020

2021/2022Technology

FAST spacecraft

Laser transmitter

Solid state laser diode

Pointing & tracking

Surface receiver

CIGS PV cells

CIGS PV array

Integration

TRL 6

TRL 6

TRL 6

TRL 3

TRL 3

TRL 4

TRL 4 TRL 6

TRL 6

TRL 7

2023/2024

TRL 3

80 % efficiency QE at 830nm/ 5 to 10 year life

0.05° pointing error / 5 to 10 year life

50 % efficiency / 5 to 10 year life

40 W/kg / 5 to 10 year life

Roadmap to achieve laser power beaming technology by 2018Roadmap to achieve laser power beaming technology by 2018

47

Task IIOther Surface Energy Storage Technologies

for Comparison

48

Surface Energy Storage Technology Trade Space

Li-S (Sion)LiTE*STARNano Safe

* Reference Technology: Fission Surface Power

Regenerative Fuel Battery

Separated

Flywheel

Li-IonUnitized

Lunar Surface Energy Storage*

Supercapacitor

Pressured Storage Cryogenic Storage

Thermal Storage

Thin Film

EEStor

Integrated with Propellant Tanks

SeparatedTanks

08PD-167-003

Power (to user) kWe 5

Moon Sidereal Periods hours 708.7

Lunar Eclipse hours 354.0

Lunar Sunlit hours 354.7

Total ES kW-h 2,000

Lunar Night Sink T K 100

Lunar Day Sink (with Apron) K 250

Solar Insolation W/m2 1,367

Trade Case and

Assumptions

Regenerative Fuel Cells

Pressurized Storage

49

Regone Plot of Energy Storage Technologies

Regenerative fuel cell (RFC) is suitable for long duration energy storageRegenerative fuel cell (RFC) is suitable for long duration energy storage

0.001

0.01

0.1

1

10

100

1000

10000

0.001 0.01 0.1 1 10 100 1000

Specific Power (kW/kg)

Spec

ific

Ener

gy (W

h/kg

)100 hrs 1 hr 1 minute

1 second

10 ms

Capacitor

EEStorLi-S

NanoSafeRegen. Fuel Cells

Adv. Battery

Flywheel

SupercapacitorFlow Battery

50

RFC with High Pressurized and Cryogenic Reactant Storage

Schematic of RFC with pressurized tank storage Schematic of RFC with cryogenic storage

H2H2

State of the Art Advanced

PEM FC PEM ElectrolysisH P Storage

AFCPEM ElectrolysisH P Storage

Unitized Regen. Fuel Cell (URFC)H P Storage

Advanced URFCCryogenic Storage

AFC or HE PEMFCPEM ElectrolysisCryogenic Storage

FC Efficiency (%) @BOL 54 65 50 54 70

Electrolysis Efficiency (%) 93 93 90 93 93

SOTA AFC provides highest FC efficiency & PEMFC needs to improve efficiencySOTA AFC provides highest FC efficiency & PEMFC needs to improve efficiency

51

Various RFC Configurations CBE Mass Data & TRL

PEM FC PEM Electrolysis

H P Storage

AFCPEM Electrolysis

H P Storage

Unitized Regen. Fuel Cell (URFC)

H P StorageAdvanced URFC

Cryogenic StorageAFC PEM Electrolysis

Cryogenic StorageFC Efficiency (%) @BOL 54 65 50 54 70Electrolysis Efficiency (%) 93 93 90 93 93RFC RT Efficiency (%) 50 60 45 50 65Power for Charging (kWe) based on power input 10.5 8.7 11.7 10.5 8.1FC Specific Power (kW/kg) based on power output 0.167 0.102 0.102Electrolysis Specific Power (kW/kg) based on power for charging 0.75 0.75 0.75RFC Specific Power (kW/kg) based on power output 0.116 0.083 0.0565 0.09825 0.084FC Operating T (oC) 80 80 90 120 120Reactant Usable (%) 95 95 95 95 95Working Performance Factor for Hydrogen Tank (379330 psi-in3/lb) 379330 379330 379330Working Performance Factor for Oxygen Tank (379330 psi-in3/lb) 379330 379330 379330controls, and instrumentation) (kW/kg) based on power output 0.06 0.06 0.06 0.05 0.05

Mass (kg) Mass (kg) Mass (kg) Mass (kg) Mass (kg)RFC 46 63 93 54 62Radiator 15 10 16 10 5Reactant (H2) 112 93 125 112 87Reactant (O2) 897 745 1001 897 692Tanks 2732 2312 3238 467 393Structure, ancillary components, piping, controls, and instrumentation 83 83 83 125 108PMAD 10 10 10 10 10Drying/Liquification equipment 132 104Power for cryogenic storage 338 267Others (additional radiator, piping) 19 10Additional Solar Array Power for Energy Storage 31 26 35 22 17Additional PMAD 5 4 6 5 4

TOTAL Mass 3931 3347 4607 2191 1760Specific Energy (W-h/kg), based on energy output 509 598 434 913 1137

TRL > 5 > 5 5 3 3

Regenerative Fuel CellState of the Art Advanced

Regenerative Fuel Cell with Cryogenic Storage has potential to provide > 1000 W-hr/kg for Lunar surface energy storage Regenerative Fuel Cell with Cryogenic Storage has potential to provide > 1000 W-hr/kg for Lunar surface energy storage

AFC or HE PEMFC

52

Li Based Battery Mass Data

Specific Energy Cycles(W-hr/kg) (#)

Li Based Battery (SOTA)Li-Ion > 200 > 2,000 (DOD of 80%)

Li Based Battery (Advanced, by 2015)

Thin Film Li Sulfur (Sion Power) > 300Currently low cycles

Improvement expectedLiTE*STAR Thin film Li-Polymer Solid State (Infinite Power Solutions/ORNL) 300 70,000

Type

• Li battery technology has been advanced quickly: 3X in 10 years• Estimated at >300 W-hr/kg by 2015, Sion Power’s Li-S battery is a

promising battery technology,• (SOTA) Li-S suffers low efficiency (80% vs. >98% of Li-ion battery)

and low cycle life

53

Li-Ion Battery FlywheelLi Polymer Battery*

Thin Film Li Sulfur(Sion Power)

Battery-Supercapacitor Hybrid EEStor

Thermal Storage FSP*

RT Efficiency (%) @ BOL 95 92 99 80 98DOD (%) 80 80 90 90 95Power for Charging (kWe) based on power input 6.9 7.1 5.9 7.3 5.6Battery (kW-hr/kg) 0.150 0.100 0.240 0.320 0.224Operating T (oC) 40 60 120 65 40Structure, controls \and instrumentation (kW/kg) 0.02 0.03 0.02 0.02 0.02

Mass (kg) Mass (kg) Mass (kg) Mass (kg) Mass (kg)Energy Storage 17544 27174 9353 8681 9590Radiator 2 2 2 2 2PMAD 10 10 10 10 10Structure, ancillary components, piping, controls, and instrumen 250 166.6666667 250 250 250

Additional Solar Array Power for Energy Storage 20 21 17 22 17Additional PMAD 3 4 3 4 3

TOTAL Mass 17830 27377 9635 8968 9872 18460 < 3,000Specific Energy (W-h/kg), based on energy output 112 73 208 223 203 108 > 600

TRL > 6 > 6 5 4 4 > 6 4 to 5

State of the Art Advanced

Task 2 CBE Mass Data & TRL for Surface ES Technologies (other than FC based)

• Note: mass data is extrapolated from a 40 kWe FSP.

Advanced (TRL>6 between 2015 to 2018) energy storage technologies, other than RFC, can achieve ~200 W-hr/kg Advanced (TRL>6 between 2015 to 2018) energy storage technologies, other than RFC, can achieve ~200 W-hr/kg

54

Summary of Task II Energy Storage Technology Trade Results

• RFC with cryogenic storage is a preferred lunar surface energy storage technology• AFC or high efficiency PEMFC is recommended as the fuel cell technology

• AFC suffers life issue• High efficiency PEMFC development is required: from 55% (SOTA) to 70%

• Cryogenic storage of RFC reactants reduces mass by a half and volume by 4X comparing with high pressure storage RFC configuration

• Waste heat from relative less efficient RFC can be integrated with ECLSS• Due to its limited applications, the RFC technology may require more funding to

develop in order to achieve TRL > 6 by 2015

RFC with Cryogenic Storage is preferred for lunar surface energy storage RFC with Cryogenic Storage is preferred for lunar surface energy storage

Lunar Surface Energy Storage

TechnologiesSpecific

Energy (CBE - W-hr/kg)

System Efficiency, Not Including Waste Heat Application (%)

Energy Density (W-hr/l) / Volume

Failure Tolerance

Risk (TRL)

RFC (AFC) w Pressurized Storage 705 60 > 200 Medium > 5Li-Ion 112 95 < 350 High > 6

Flywheel 73 92 < 200 Medium > 6RFC (AFC) w Cryogenic Storage 1153 65 > 1000 Medium 3

Li-Polymer 208 99 < 600 High 5Thin Film Li-S Battery 223 80 < 600 High 4

EEStor 203 98 > 1000 High 4

Advanced

Technical Performance

SOTA

55

Summary of Power Beaming vs. Ad. RFC w/ Cryogenic Storage

Lunar Surface Energy Storage

Technical Performance Programmatic

Technology ElementMass (CBE)(kg)

Stowed Volume (W-hr/l)

Oper. Effect.

(mobility)

Specific Energy (CBE)

(W-hr/kg)

* Sys. Eff. (%)

Redund-

ancy

Risk (TRL)

Rel.Dev. Cost

Integration with Other Systems

Space 1,000 N/A High N/A 30 x 50 Low 3 Med. Telecom.

Surface 3,997 720 Med. 500 50 High 3 Low

Space 2,000 N/A High N/A 30 x 50 High 3 Med. Telecom.

Surface 3,997 720 Med. 500 50 High 3 Low

RFC (cryogenic storage) Surface 1,735 > 1,000 Low 1,153 30 x 65 Med. 3 Med.

ECLSS, Ascent

Propellant, ISRU

Power Beaming (two spacecrafts)

Power Beaming (one spacecraft)

* Spacecraft power system (PV array) and laser; RFC power system (PV array) and fuel cell, respectively

• From mass point of view, power beaming is no better than RFC technology• Power Beaming has advantages of mobility & redundancy (for 2 s/c case)• Power Beaming and RFC w/ cryo. storage are at low TRL (3: proof of concept)

• Expect both require relative medium development cost comparing with FSP• Both are highly “integrate-able” with other subsystems for optimal application

56

• Laser power beaming provides a feasible option for lunar energy storage• Key benefits include added value of mobility and integration of telecom. • A two spacecraft configuration is recommended due to redundancy• A near optimal configuration includes orbital period of 16.1 hours which results

in a “gap” in coverage of less than 16 hours• A surface energy storage on the order of 525 kW-hr, for 5 kWe continuous

power, which reduced requirement by > 70%• Laser power beaming & CIGS PV receiver provides the smallest footprint• Recommend developing DARPA FAST spacecraft with electric propulsion • For min. mass, power beaming is no better than Regenerative Fuel Cells

• Regenerative Fuel Cells with cryogenic reactant storage is a feasible and preferred lunar surface energy storage tech

• High eff. FC, Alkaline FC or PEM FC, is required

Study Conclusions

57

Recommendations for Forward Work• Continue study of spacecraft laser power beaming technologies

• Power generation, PMAD, laser beam generation and transmission, thermal management for the laser subsystem, and laser beam pointing and tracking• Investigate alternative orbit strategies (than “frozen orbit”), such as matching the precession rates of the moon, to maximize coverage time

• Conceptual synthesis of laser telecom with laser power beaming• Perform a subscale laser photovoltaic ground demonstration

• Demonstration performed in a simulated space environment and include the 830 nm solid state laser diode and CIGS photovoltaic array receiver subsystem technologies

• Demonstrate Regenerative Fuel Cells with cryogenic reactants storage• Optimal assembly, more reliable & failure tolerance, to get TRL > 6 by 2018

• Initiate an integrated study for resources of “Just-in-time resource management”

• Study of optimal usage of limited resources for Altair, Habitation/ECLSS, Energy Storage, Ascent Propellants, ISRU

• Six essential compounds: H2, O2, H2O, CH4, CO2 and CO

Page 58Hamilton Sundstrand Rocketdyne Proprietary 58

59

Backup

60

Libration Points in the Earth-Moon System

Earth Moon

L3

L4

L5

L2L1

R

d1 d2

Mean Distances

R

d1

d2

384,000 km

58,000 km

64,500 km

61

Representative Earth Lunar Libration Point Halo Orbit

Moon

Moon’s orbitHalo orbit

L2

Moon

Halo orbit

Power spacecraft

62

Total Mass – 2 to 5 kWe Surface Power

0

10

20

30

40

50

60

1 Sat., L1Halo

1 Sat., L2Halo

1 Sat., Polar 1 Sat.,Equatorial

1 Sat., 45°Inclination

2 Sat., 45°Inclination

ConventionalSpacecraft

EnergyStorage

Architecture Concept

Mas

s (t)

10, 14 and 22 kWe spacecraft power

SurfacePower, kWe

2 3 5

63

Surface Element Mass – 2 to 5 kWe Surface Power

0

10

20

30

40

50

60

1 Sat., L1Halo

1 Sat., L2Halo

1 Sat., Polar 1 Sat.,Equatorial

1 Sat., 45°Inclination

2 Sat., 45°Inclination

ConventionalSpacecraft

EnergyStorage

Architecture Concept

Surf

ace

Elem

ent M

ass

(t)

10, 14 and 22 kWe spacecraft power

SurfacePower, kWe

2 3 5

64

Relative Cost Comparison – 2 to 5 kWe Surface Power

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

1 Sat., L1Halo

1 Sat., L2Halo

1 Sat., Polar 1 Sat.,Equatorial

1 Sat., 45°Inclination

2 Sat., 45°Inclination

ConventionalSpacecraft

EnergyStorage

Architecture Concept

Rel

ativ

e C

ost

SurfacePower, kWe

2 3 5

10, 14 and 22 kWe spacecraft power

65

System Efficiency – 2 to 5 kWe Surface Power

0

10

20

30

40

50

60

70

80

90

100

1 Sat., L1Halo

1 Sat., L2Halo

1 Sat., Polar 1 Sat.,Equatorial

1 Sat., 45°Inclination

2 Sat., 45°Inclination

ConventionalSpacecraft

EnergyStorage

Architecture Concept

Syst

em E

ffici

ency

(per

cent

)

SurfacePower, kWe

2 3 5

10, 14 and 22 kWe spacecraft power

Combined Space and Surface Elements

66

System Efficiency – 2 to 5 kWe Surface Power

0

10

20

30

40

50

60

70

80

90

100

1 Sat., L1Halo

1 Sat., L2Halo

1 Sat., Polar 1 Sat.,Equatorial

1 Sat., 45°Inclination

2 Sat., 45°Inclination

ConventionalSpacecraft

EnergyStorage

Architecture Concept

Syst

em E

ffici

ency

(per

cent

)

Surface Element Only

SurfacePower, kWe

2 3 5

10, 14 and 22 kWe spacecraft power

67

Propellant Mass – 2 to 5 kWe Surface Power

0

5

10

15

20

25

30

35

40

45

50

1 Sat., L1 Halo 1 Sat., L2 Halo 1 Sat., Polar 1 Sat.,Equatorial

1 Sat., 45°Inclination

2 Sat., 45°Inclination

ConventionalSpacecraft

Architecture Concept

Stat

ion-

keep

ing

Prop

ella

nt M

ass

(kg)

10, 14 and 22 kWe spacecraft power

SurfacePower, kWe

2 35

68

Candidate Orbits

Alt p Alt a R p R a a e V p V a Period Inckm km km km km - km/s km/s hours deg

11,039 11,039 12,777 12,777 12,777 0.000000 0.619 0.619 36.000 0.0008,013 8,013 9,751 9,751 9,751 0.000000 0.709 0.709 24.000 0.0006,311 6,311 8,049 8,049 8,049 0.000000 0.780 0.780 18.000 0.0005,928 16,149 7,666 17,887 12,777 0.400001 0.946 0.406 35.998 45.0004,112 11,913 5,850 13,651 9,751 0.400000 1.083 0.464 24.000 45.0003,091 9,531 4,829 11,269 8,049 0.400000 1.192 0.511 18.000 45.000

Alt p Alt a R p R a a e V p V a Period Inckm km km km km - km/s km/s hours deg

11,039 11,039 12,777 12,777 12,777 0.000000 0.619 0.619 36.000 0.0008,013 8,013 9,751 9,751 9,751 0.000000 0.709 0.709 24.000 0.0006,311 6,311 8,049 8,049 8,049 0.000000 0.780 0.780 18.000 0.0005,928 16,149 7,666 17,887 12,777 0.400001 0.946 0.406 35.998 45.0004,112 11,913 5,850 13,651 9,751 0.400000 1.083 0.464 24.000 45.0003,091 9,531 4,829 11,269 8,049 0.400000 1.192 0.511 18.000 45.000

Alt p Alt a R p R a a e V p V a Period Inckm km km km km - km/s km/s hours deg

11,041 11,041 12,777 12,777 12,777 0.000000 0.619 0.619 36.000 90.0008,015 8,015 9,751 9,751 9,751 0.000000 0.709 0.709 24.000 90.0006,313 6,313 8,049 8,049 8,049 0.000000 0.780 0.780 18.000 90.000

Alt p Alt a R p R a a e V p V a Period Inckm km km km km - km/s km/s hours deg

11,041 11,041 12,777 12,777 12,777 0.000000 0.619 0.619 36.000 90.0008,015 8,015 9,751 9,751 9,751 0.000000 0.709 0.709 24.000 90.0006,313 6,313 8,049 8,049 8,049 0.000000 0.780 0.780 18.000 90.000

a semi-major axis Alta periapsis altitudee eccentricity Altp apoapsis altitude

Vp periapsis velocity Rp periapsis radiusVa apoapsis velocity Ra apoapsis radius

Period orbit period Inc inclination

a semi-major axis Alta periapsis altitudee eccentricity Altp apoapsis altitude

Vp periapsis velocity Rp periapsis radiusVa apoapsis velocity Ra apoapsis radius

Period orbit period Inc inclination

69

Candidate Frozen Orbits

Alt p Alt a R p R a a e V p V a Period Inckm km km km km - km/s km/s hours deg

1,897 6,912 3,635 8,650 6,143 0.4082 1.378 0.579 12.000 45.0003,025 9,597 4,763 11,335 8,049 0.4082 1.204 0.506 18.000 45.0004,032 11,993 5,770 13,731 9,751 0.4082 1.094 0.460 24.000 45.000

719 8,090 2,457 9,828 6,143 0.6000 1.787 0.447 12.000 51.7071,482 11,140 3,220 12,878 8,049 0.6000 1.561 0.390 18.000 51.7072,162 13,863 3,900 15,601 9,751 0.6000 1.418 0.355 24.000 51.707

Alt p Alt a R p R a a e V p V a Period Inckm km km km km - km/s km/s hours deg

1,897 6,912 3,635 8,650 6,143 0.4082 1.378 0.579 12.000 45.0003,025 9,597 4,763 11,335 8,049 0.4082 1.204 0.506 18.000 45.0004,032 11,993 5,770 13,731 9,751 0.4082 1.094 0.460 24.000 45.000

719 8,090 2,457 9,828 6,143 0.6000 1.787 0.447 12.000 51.7071,482 11,140 3,220 12,878 8,049 0.6000 1.561 0.390 18.000 51.7072,162 13,863 3,900 15,601 9,751 0.6000 1.418 0.355 24.000 51.707

a semi-major axis Alta periapsis altitudee eccentricity Altp apoapsis altitude

Vp periapsis velocity Rp periapsis radiusVa apoapsis velocity Ra apoapsis radius

Period orbit period Inc inclination

a semi-major axis Alta periapsis altitudee eccentricity Altp apoapsis altitude

Vp periapsis velocity Rp periapsis radiusVa apoapsis velocity Ra apoapsis radius

Period orbit period Inc inclination

70

Representative STK Gaps Chart – 3 spacecraft Case

Previous Lunar Night Cycle Next Lunar Night Cycle

Lunar Day Cycle with Earth Eclipse

STK v8.1.3 simulation results. 10° latitude, -45° longitude surface site.

71

AGI spacecraft Tool Kit (STK) Simulation Constraints

• Unconstrained lighting– Imposed only spacecraft-to-surface line-of-sight constraint

A direct line-of-sight must exist between surface site and spacecraft for valid access

• Constrained lighting– Imposed spacecraft-to-surface line-of-sight constraint– Imposed spacecraft sunlight and surface darkness constraints

spacecraft must be in direct sunlight and either umbra or penumbra condition must exist at surface site for valid access

• Lunar surface sites arbitrarily selected to maximize coverage– Initial site and revised surface sites

45° latitude, 60° longitude, 0 km altitude (initially)10° latitude, -45° longitude, 0 km altitude (revised)

• Lunar gravity model and orbit propagator– Initially used simple lunar gravity model & low precision orbit propagator

Later analysis used high precision orbit propagator (HPOP) with permanent tides– Lunar gravity third bodies included Earth and Sun.

72

Access Summary – 18 hour Orbits

0

10

20

30

40

50

60

70

80

90

100

0 50 100 150 200 250 300 350 400 450 500

Access Number (01 Jan through 31 Dec 2020)

Sate

llite

Acc

ess

Tim

e (p

erce

nt o

f orb

ital p

erio

d)18 hour circular orbit 18 hour elliptical orbit18 hour elliptical orbit (constraints) 18 hour circular orbit (constraints)

STK v8.1.3 simulation results. 45° latitude, 60° longitude surface site.

73

Coverage Summary – 18 hour Orbits

0

10

20

30

40

50

60

70

80

90

100

Jan Feb Mar Apr May Jun Jul Aug Sep Oct Nov Dec

Month (2020)

Cov

erag

e Ti

me

(per

cent

of L

unar

dar

knes

s pe

riod)

18 hour elliptical orbit (constraints) 18 hour circular orbit (constraints)

STK v8.1.3 simulation results, 45° inclination orbit. 45° latitude, 60° longitude surface site.

74

Coverage Summary (Jan 2020, 45° Orbit Inclination)

0

10

20

30

40

50

60

70

80

90

100

8 12 18 24 36 48

Orbital Period (hours)

Cov

erag

e Ti

me

(per

cent

of L

unar

dar

knes

s pe

riod)

Eccentricity = 0.7 Eccentricity = 0.6 Eccentricity = 0.4 Eccentricity = 0.0

10° latitude, -45° longitude surface site.STK v8.1.3 simulation results, 45° inclination orbit.

75

Coverage Summary (Jan 2020, 0° Orbit Inclination)

0

10

20

30

40

50

60

70

80

90

100

8 12 18 24 36 48

Orbital Period (hours)

Cov

erag

e Ti

me

(per

cent

of L

unar

dar

knes

s pe

riod)

Eccentricity = 0.7 Eccentricity = 0.6 Eccentricity = 0.4 Eccentricity = 0.0

STK v8.1.3 simulation results. 10° latitude, -45° longitude surface site.

76

Maximum Range (Jan 2020, 48 hour Orbital Period)

STK v8.1.3 simulation results.

0

5,000

10,000

15,000

20,000

25,000

30,000

0.7 0.6 0.4 0.0

Orbital Eccentricity (none)

Max

imum

Sat

ellit

e-to

-Lun

ar S

ite R

ange

(km

)

0 deg Inclination 45 deg Inclination

10° latitude, -45° longitude surface site.

77

Spacecraft Orbit and Lunar Surface Site AssumptionsSurface Site

Lat (deg) Lon (deg) Alt. (km)10 -45 0

SatellitePeriod (hrs) SMA (km) Eccentricity Inc. (deg) Arg. (deg) RAAN (deg) TA (deg)

24 8,049 0.6 45 300 0 120 01 Jan 2020 12:00:00 AMEpoch (UTCG)

78

Access Summary – 24 hour Elliptical Orbit Period

0

10

20

30

40

50

60

70

80

90

100

0 50 100 150 200 250 300

Access Number (01 Jan through 31 Dec 2020)

Sate

llite

Acc

ess

Tim

e (p

erce

nt o

f orb

ital p

erio

d)

STK v8.1.3 simulation results. 10° latitude, -45° longitude surface site.

79

Coverage Summary – 24 hour Elliptical Orbit Period

0

10

20

30

40

50

60

70

80

90

100

Jan Feb Mar Apr May Jun Jul Aug Sep Oct Nov Dec

Month (2020)

Cov

erag

e Ti

me

(per

cent

of L

unar

dar

knes

s pe

riod)

10° latitude, -45° longitude surface site.STK v8.1.3 simulation results.

80

Coverage Details – 24 hour Elliptical Orbit Period

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21

Access Number (01 Jan through 05 Feb 2020)

Cov

erag

e Ti

me

(day

s)

10° latitude, -45° longitude surface site.STK v8.1.3 simulation results.

Previous Night Cycle Next Lunar Night Cycle

Lunar Day Cycle with

Earth Eclipse

81

Receiver Length Variation with Pointing Error

0

20

40

60

80

100

120

140

0.00 0.05 0.10 0.15 0.20

Laser Beam Pointing Error (deg)

Luna

r Sur

face

Col

lect

or L

engt

h (m

)

(64 m at 0.05° pointing error, 30 ° minimum elevation angle and 9,977 km max. range)

10° latitude, -45° longitude surface site.Elliptical orbit, 45° inclination, 16.1 hour orbit period.

82

Receiver Width Variation with Pointing Error

0

20

40

60

80

100

120

0.00 0.05 0.10 0.15 0.20

Laser Beam Pointing Error (deg)

Luna

r Sur

face

Col

lect

or W

idth

(m)

(41 m at 0.05° pointing error, 30 ° minimum elevation angle and 9,977 km range)

10° latitude, -45° longitude surface site.Elliptical orbit, 45° inclination, 16.1 hour orbit period.

83

Surface Receiver Mass and Stowed Volume

Module mass includes 2 kg for each of two “tent” sections (4 kg total per module).

Deployed length is for active receiver only, and does not include “tent” section length (3.4 m each end).

Deployed Module Pwr Cable TotalLength Width Area Mass Mass Modules Mass

m m m2 kg kg kg24.0 2.5 60 34 1.3 2 7140.0 2.5 100 54 2.1 2 11248.0 2.5 120 64 2.5 2 13354.0 2.5 135 72 2.8 2 14958.0 2.5 145 77 3.0 2 15962.0 2.5 155 82 3.2 2 16964.1 2.5 160 84 3.3 4 350

Total 1,143

Deployed Module Pwr Cable TotalLength Width Area Mass Mass Modules Mass

m m m2 kg kg kg24.0 2.5 60 34 1.3 2 7140.0 2.5 100 54 2.1 2 11248.0 2.5 120 64 2.5 2 13354.0 2.5 135 72 2.8 2 14958.0 2.5 145 77 3.0 2 15962.0 2.5 155 82 3.2 2 16964.1 2.5 160 84 3.3 4 350

Total 1,143

Deployed Number Stowed Stowed Stowed Number TotalLength of Folds Length Height Width Mod. Vol. of Modules Volume

m m m m m3 m3

24.0 24 1.0 0.057 2.5 0.143 2 0.340.0 40 1.0 0.061 2.5 0.153 2 0.348.0 48 1.0 0.063 2.5 0.158 2 0.354.0 40 1.4 0.061 2.5 0.207 2 0.458.0 58 1.0 0.066 2.5 0.164 2 0.362.0 62 1.0 0.067 2.5 0.167 2 0.364.1 64 1.0 0.067 2.5 0.169 4 0.7

Total 2.7

Deployed Number Stowed Stowed Stowed Number TotalLength of Folds Length Height Width Mod. Vol. of Modules Volume

m m m m m3 m3

24.0 24 1.0 0.057 2.5 0.143 2 0.340.0 40 1.0 0.061 2.5 0.153 2 0.348.0 48 1.0 0.063 2.5 0.158 2 0.354.0 40 1.4 0.061 2.5 0.207 2 0.458.0 58 1.0 0.066 2.5 0.164 2 0.362.0 62 1.0 0.067 2.5 0.167 2 0.364.1 64 1.0 0.067 2.5 0.169 4 0.7

Total 2.7

1,143 kg and 2.7 m3 surface receiver, 5 kWe power beaming architecture, CIGS PV technology, Jan – Feb 2020.1,143 kg and 2.7 m3 surface receiver, 5 kWe power beaming architecture, CIGS PV technology, Jan – Feb 2020.

84

Northrop-Grumman FIRESTRIKE Laser

15 kW FIRESTRIKE Laser Line Replaceable Unit (LRU)

• High power, high efficiency laser technology exists• ONR Lasercomm is emerging• High power, high efficiency laser technology exists• ONR Lasercomm is emerging

85

Two-Axis Ion Thruster Gimbal

• Alliance Spacesystems developed two-axis ion thruster gimbal for Dawn asteroid mission

– 3.5 kg gimbal supports 7 kg engine– Improvement over 17 kg flight-proven

Deep Space-1 ion thruster gimbal– Two rotary actuators drive composite

struts with rod end-bearings mounted in a hexapod configuration

– Provides two-axis thrust vector pointing

A similar configuration may provide the laser subsystem “rough” pointing functionA similar configuration may provide the laser subsystem “rough” pointing function

86

Office of Naval Research Solid-State Laser

ONR Solid-State Fiber Laser ONR Lasercomm

87

Comparison Between Shuttle AFC and Lunar ES Requirement

Orbiter Alkaline FC Lunar Energy StoragePower (kWe) 12 2 to 5Duration (hr) 384* 354Energy (kW-hr) 4,694 2,000

* based on 12 kW, 70% efficiency, 5 cryogenic storage tanks (for Hydrogen and Oxygen), 95.5% Hydrogen usable and 90% Oxygen usable

• With ~3000 kg mass the Shuttle FC power can provide more than two lunar night cycles (2 months) power needs at 5 kWe

88

Li Based Battery Technical Data

Voltage Specific Energy

Energy Density

Specific Power Efficiency $/E Discharge Cycles Life

(V) (W-hr/kg) (W-hr/L) (W/kg) (%) ($/W-hr) (%/mo) (#) (years)Ni NiMH 1.2 30-80 140-300 250-1,000 66% 0.73 20% 1,000 hybrid cars 1983

Li-Ion(SOTA - Varta) 3.6 160 270 1,800 99.9% 0.4 - 0.2 5%-10% 1,200 2 to 3

best specific power: 220 W-hr/kg & density: 400 W-hr/ltheoretical >0.75kW-hr/kg & >1.8kW-hr/l 1990

Very H P (VL12V) Li-Ion (SOTA - Saft) 3.6 75 175

5,000 (18 sec at 2.5 V) 300,000 (3% DOD) 15 yrs -30 to 60 oC (discharge) 2007

Li-Ion for EV(SOTA - Yardney) 145 358 2,100 (80% DOD) -40 to 65 oC (discharge)Li-Ion Polymer(SOTA - GRC) 3.7 < 200 300 2,800 99.8% 0.4 - 0.2 5%/month >1,000 cycles 2 to 3 PMA 1996Li-Ion Polymer(SOTA - Danionics) 180 (G3) 350 (G3) 400 (80% DOD) -20 - 60 oC 2002

Thin Film Li Sulfur(Sion Power) 2.1

350 (cell) & 260 (pack)

600 (< 2015)360

600 (< 2015) 1,000 low cycles theoretical >2.5kW-hr/kg & >2.6kW- 1994

NanoSafeNano Titanate (Altairnano)

13.8 100 4,000+ 87-95% 2.5 20,000 20

increasing effective area; wide temp range (-50 to 75 oC); charge in <10 min. 2007

LiTE*STARThin film Li-Polymer Solid State (Infinite Power Solutions/ORNL)

> 3.6200

300 (< 2015)450

900 (< 2015) 6,00060,000

70,000 (< 2015)

LiPON electrolyte/separator; Li as anode; 15 μm; wide temp range (-20 to 140 oC); credit card size commercially available 2007

LiVO2 (A)/LiCoO2 (C)(Subaru ) 745Li2FePO4F(U. Waterloo)

Long (no cathode volume changes)

Utrathin LIB(MIT nano tube anode) 3XNano electrodes(France) several X several XSi nanowires on s.s. Anode(Stanford) several XLi Nickel Cathode(Tiax - Johnson Control) 40% better

Li

SinceApplication/CommentsTech. Type

89

18650 Cylindrical Li-Ion Battery Cell Capability Improvement

0

100

200

300

400

500

600

700

800

1992

1993

1994

1995

1996

1997

1998

1999

2000

2001

2002

2003

2004

2005

Future

Year

1865

0 S

peci

fic E

nerg

y (W

-hr/k

g) o

r Ene

rgy

Den

sity

(W-h

r/l)

Specific Energy

Energy Density

90

Flywheel Mass Data

65 kW-hr Flywheel (A Pair) Flywheel (SOTA)

Max. Speed (RPM x 1,000) 65

Efficiency (%) 85

Mass (kg)

Flywheel Rotor Mass 282

Motor/Generator Mass 4

Shaft Mass 2

Magnetic Bearing Mass 27

Containment Mass 366

Total 681

Specific Energy (W-hr/kg) 95

• By 2015, the estimated flywheel specific energy may still be < 150 W-hr/kg

91

Super-Capacitor Technical Data

Specific Power Specific Energy Power

DensityEnergy Density

Specific Surface Area

(W/kg) (W-hr/kg) (W/l) (W-hr/l) (m2/g)SOTA Maxwell Ultracapacitor 15000 up to 6 9 2.7V/cell

Tartu tech. (Skeleton Nanolab)

Carbide Derived Carbon 20000 8 25 13 5

high sp. area (400 - 2000 m2/kg) 400 - 2000

MIT LEES Carbon Nanotube 30 - 60 3

PowerStor Carbon Aerogel 10000 109 ("AA" size)

2.5Vhigh sp. area (400 - 1000 m2/kg) 400 - 1000

Reticle Carbon

Reticle Carbon (solid active carbon) 7500 (theoretical) 2

high sp. area (> 1000 m2/kg) > 1000

EEstor

Battery-Ultracapacitor Hybrid Barium Titanate 280-342 5

high voltage (3500V)low price ($0.04/Wh)

Adv

ance

d

OthersCompany Technology TRL

• EEStor, of Cedar Park, Texas, employing a ceramic super-capacitor with a barium-titanate insulator to achieve high specific power: 280 W-hr/kg